The occurrence of supersonic flow over a portion of a rotating compressor or fan blade is a phenomenon which is familiar to those skilled in the art of aircraft propulsion system design. Relative flow velocity is a function of the engine nacelle or inlet through flow velocity, the angular velocity of the rotating blade, and the radial distance between the blade region under consideration and the axis of rotation.
Prior art blade designers have recognized the flow disruption which results from such relative supersonic flow, and, in particular with respect to noise generation, have specified blade leading and trailing edge shapes responsive to the supersonic flow conditions which attempt to minimize the occurrence of sonic shock waves in the vicinity thereof. One such blade design is shown in U.S. Pat. No. 3,989,406 issued Nov. 2, 1976 to Bliss. The Bliss reference shows a rotor blade wherein the leading edge of that portion of the blade subject to supersonic relative airflow velocities is swept axially forward or rearward behind a Mach cone defined at each point along the blade leading edge. This critical skewing or sweeping of the blade leading edge results in the component of the relative airflow velocity normal to the leading edge of the blade having a Mach number less than 1 and therefore falling in the subsonic flow regime. The Bliss design is intended to prevent leading edge shock waves from forming thereby reducing shock related noise generation by the rotating blades.
Bliss also discusses the problem of the formation of pressure waves adjacent the shroud wall surrounding the rotating blades. These shroud wall pressure waves interact with the airflow over the blade tips and can result in a second shock wave in the vicinity of the rotating blade tips. Bliss further shows a method and means for reducing such shroud-tip interaction by contouring the shroud wall along the natural streamline deflection over the suction surface of the blade tip.
A related blade design is disclosed in U.S. Pat. No. 4,012,172 issued Mar. 15, 1977 to Schwaar et al. The Schwaar reference shows a noise reducing fan blade design wherein the leading edge of the blade is swept forwardly from the blade hub to a point at the blade midspan and then swept rearwardly to the blade tip. The leading edge sweep is, for that portion of the blade subject to supersonic relative airflow velocities, shaped to fall within the Mach cone of the upstream adjacent leading edge points, thus achieving the subsonic normal velocity component indicated in the Bliss reference. Schwaar also describes a method for configuring such blades to minimize internal blade bending and attachment stresses by balancing the centers of gravity of the successive blade transverse segments about the attachment radius such that radial forces induced by the movement of the blade about the rotation axis are substantially balanced.
Finally, U.S. Pat. No. 4,358,246 issued Nov. 7, 1982 to Hanson et al also shows the contouring of a supersonic prop fan blade so that both the leading and trailing edges of the blade are swept behind their corresponding Mach surfaces. The pressure spike induced in prior art blades by the trailing wave is thus minimized along with the noise generation resulting therefrom.
The prior art blade designs thus discussed have been directed toward minimizing the noise generation which can occur in the case of improperly designed supersonic fan blades. Such noise reducing designs as are present in the prior art, while effective in reducing leading and trailing edge shock waves, do not address the occurrence of a strong compression wave front at the suction surface of a supersonic blade intermediate the leading and trailing edges. The compression wave front is the result of the recompression of the air stream flowing over the suction surface of the fan blade as the local airflow ceases its relative acceleration with respect to the blade surface and begins to decelerate adjacent that portion of the blade which lies downstream of the point of maximum blade camber.
As airflow relative velocity decreases, the static air pressure rises from the minimum value which occurs coincident with the maximum relative airflow velocity. This increase in pressure forms a pressure wave which, reinforced by the pressure waves of adjacent points along the blade suction surface, contributes to the creation of a rapid static pressure jump wherein the pressure gradient at the blade surface rises too rapidly to permit smooth flow in the boundary layer downstream thereof. This strong compression wave front induced by the reinforced compression waves at the surface of the blade detaches and disrupts the downstream blade surface boundary layer, resulting in flow recirculation and other irreversible losses which diminish overall blade performance by both increasing blade drag and decreasing the amount of airflow turning or static pressure rise achieved.
Prior art supersonic blade designs have attempted to delay such separation and its deleterious performance effects by shaping the transverse cross section of individual blade segments such that the point of maximum blade camber is located toward the trailing edge of the blade, with the exact location for each blade section based on prior state of the art unswept blade cascade test data correlations. The separation of the boundary layer which occurs due to the recompression of the flowing air downstream of the maximum blade camber point is thus able to act on only a small portion of the blade suction surface.
The prior art technique, while effective to a degree, does not eliminate the disruption of the boundary layer and the resulting inefficiencies and irreversibilities, but only reduces the total negative effect thereof. What is required is a blade design which weakens or eliminates the compression wave front to such a degree that separation of the airflow boundary layer at the suction surface of the rotor blade does not occur.